With reference to FIG. 1, a ducted fan gas turbine engine is generally indicated at 10 and has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, an intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
During operation, air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
The compressors each comprise a number of rotor discs, each carrying a set of rotor blades having an aerofoil configuration. The discs are bolted or welded together to form a compressor drum. The rotor blades may be affixed to the discs in an axial or a circumferential fixing arrangement. Circumferential fixing is generally used in the rear stages of the compressors as it is simpler and cheaper (albeit less robust) than axial fixing.
Circumferential fixing involves machining a circumferentially-extending groove around the outer rim of each disc and then slotting the blade roots into the groove.
The circumferentially-extending groove typically has a symmetrical dove-tailed profile with multiple radii in the bulb of the dovetail to minimise stresses within the groove arising from loads applied by the blades. Minimising stresses within the groove allows a reduction in the amount and therefore weight of disc material surrounding the groove. Reduced weight leads to increased engine efficiency.
It is known to provide a bridging section between adjacent rotor discs. The bridging section provides bracing between circumferential grooves on adjacent rotor discs above the gauge plane of the rotor disc and limits distortion under the blade loads in operation. Static vanes can project from an outer casing towards the bridging sections. A spacer portion spaces adjacent rotor discs on an opposing side of the rotor disc to the bridging section.
Reducing the amount of disc material around the circumferentially-extending groove proximal the bridging section leads to a desirable weight reduction as discussed above and, furthermore, reduces stresses at the weld join between adjacent discs by reducing the thermal gradient between the weld and the rim. However, stresses are increased in the thinned area of the rotor disc.
It is a preferred aim of the present invention to provide a disc structure that can minimise the weight of the disc whilst maintaining acceptable stresses for the life of the compressor.